示例#1
0
        /// <summary>
        /// 计算绕地航天器主要的加速度。
        ///   Computes the acceleration of an Earth orbiting satellite due to
        ///    - the Earth's harmonic gravity field,
        ///    - the gravitational perturbations of the Sun and Moon
        ///    - the solar radiation pressure and
        ///    - the atmospheric drag
        /// </summary>
        /// <param name="Mjd_TT"> Mjd_TT  Terrestrial Time (Modified Julian Date)</param>
        /// <param name="satXyzIcrf"> r 卫星在国际天球参考框架的位置  Satellite position vector in the ICRF/EME2000 system</param>
        /// <param name="satVelocityIcrf"> v 卫星在国际天球参考框架的速度 Satellite velocity vector in the ICRF/EME2000 system</param>
        /// <param name="Area">  Area    Cross-section </param>
        /// <param name="mass">  mass 质量  Spacecraft mass</param>
        /// <param name="solarRadiPresCoeff"> coefOfRadiation  光压系数  Radiation pressure coefficient</param>
        /// <param name="atmDragCoefficient"> coefOfDrag  大气阻力系数  Drag coefficient</param>
        /// <returns> Acceleration (a=d^2r/dt^2) in the ICRF/EME2000 system</returns>
        public static Geo.Algorithm.Vector AccelerOfMainForces(double Mjd_TT, Geo.Algorithm.Vector satXyzIcrf, Geo.Algorithm.Vector satVelocityIcrf, double Area, double mass, double solarRadiPresCoeff, double atmDragCoefficient)
        {
            double Mjd_UT1;

            Geo.Algorithm.Vector a = new Geo.Algorithm.Vector(3), r_Sun = new Geo.Algorithm.Vector(3), r_Moon = new Geo.Algorithm.Vector(3);
            Matrix T = new Matrix(3, 3), E = new Matrix(3, 3);

            // Acceleration due to harmonic gravity field
            Mjd_UT1 = Mjd_TT;

            T = IERS.NutMatrix(Mjd_TT) * IERS.PrecessionMatrix(MJD_J2000, Mjd_TT);
            E = IERS.GreenwichHourAngleMatrix(Mjd_UT1) * T;

            a = AccelerOfHarmonicGraviFiled(satXyzIcrf, E, Grav.GM, Grav.R_ref, Grav.CS, Grav.n_max, Grav.m_max);

            // Luni-solar perturbations
            r_Sun  = CelestialUtil.Sun(Mjd_TT);
            r_Moon = CelestialUtil.Moon(Mjd_TT);

            a += AccelerOfPointMass(satXyzIcrf, r_Sun, OrbitConsts.GM_Sun);
            a += AccelerOfPointMass(satXyzIcrf, r_Moon, OrbitConsts.GM_Moon);

            // Solar radiation pressure
            a += CelestialUtil.Illumination(satXyzIcrf, r_Sun) * AccelerOfSolarRadiPressure(satXyzIcrf, r_Sun, Area, mass, solarRadiPresCoeff, OrbitConsts.PressureOfSolarRadiationPerAU, AU);

            // Atmospheric drag
            a += AccelerDragOfAtmos(Mjd_TT, satXyzIcrf, satVelocityIcrf, T, Area, mass, atmDragCoefficient);

            // Acceleration
            return(a);
        }
示例#2
0
        /// <summary>
        ///   Computes the acceleration of an Earth orbiting satellite due to
        ///    - the Earth's harmonic gravity field,
        ///    - the gravitational perturbations of the Sun and Moon
        ///    - the solar radiation pressure and
        ///    - the atmospheric drag
        /// </summary>
        /// <param name="Mjd_TT">Mjd_TT      Terrestrial Time (Modified Julian Date)</param>
        /// <param name="r">r           Satellite position vector in the ICRF/EME2000 system</param>
        /// <param name="v">v           Satellite velocity vector in the ICRF/EME2000 system</param>
        /// <param name="Area"> Area        Cross-section </param>
        /// <param name="mass"> mass        Spacecraft mass</param>
        /// <param name="coefOfRadiation">coefOfRadiation          Radiation pressure coefficient</param>
        /// <param name="coefOfDrag">coefOfDrag          Drag coefficient</param>
        /// <param name="n"></param>
        /// <param name="m"></param>
        /// <param name="isConsiderSun">是否启用太阳质量影响</param>
        /// <param name="isConsiderMoon">是否启用月亮质量影响</param>
        /// <param name="isConsiderSolarRadi">是否启用太阳光压影响</param>
        /// <param name="isConsiderDrag">是否考虑大气影响</param>
        /// <returns>Acceleration (a=d^2r/dt^2) in the ICRF/EME2000 system</returns>
        public Geo.Algorithm.Vector GetAcceleration(double Mjd_TT, Geo.Algorithm.Vector r, Geo.Algorithm.Vector v)
        {
            // Acceleration due to harmonic gravity field
            var Mjd_UT1 = Mjd_TT + (IERS.GetUT1_UTC(Mjd_TT) - IERS.GetTT_UTC(Mjd_TT)) / 86400.0;

            var T = IERS.NutMatrix(Mjd_TT) * IERS.PrecessionMatrix(OrbitConsts.MJD_J2000, Mjd_TT);
            var E = IERS.GreenwichHourAngleMatrix(Mjd_UT1) * T;

            var a = Force.AccelerOfHarmonicGraviFiled(r, E, Force.Grav.GM, Force.Grav.R_ref, Force.Grav.CS, Option.MaxDegree, Option.MaxOrder);

            // Luni-solar perturbations
            var r_Sun = CelestialUtil.Sun(Mjd_TT);

            if (Option.EnableSun)
            {
                a += Force.AccelerOfPointMass(r, r_Sun, OrbitConsts.GM_Sun);
            }
            if (Option.EnableMoon)
            {
                var r_Moon = CelestialUtil.Moon(Mjd_TT);
                a += Force.AccelerOfPointMass(r, r_Moon, OrbitConsts.GM_Moon);
            }

            // Solar radiation pressure
            if (Option.EnableSolarRadiation)
            {
                a += Force.AccelerOfSolarRadiPressure(r, r_Sun, Option.Area, Option.Mass, Option.CoefOfRadiation, OrbitConsts.PressureOfSolarRadiationPerAU, OrbitConsts.AU);
            }
            // Atmospheric drag
            if (Option.EnableDrag)
            {
                a += Force.AccelerDragOfAtmos(Mjd_TT, r, v, T, Option.Area, Option.Mass, Option.CoefOfDrag);
            }
            // Acceleration
            return(a);
        }
        static void Main0(string[] args)
        {
            // Constants

            const int    N_Step = 8;
            const double Step   = 0.5; // [d]


            // Sample Chebyshev coefficients from JPL DE405 for geocentric lunar
            // coordinates in the ITRF(EME2000) frame valid from 2006/03/14 TDB
            // to 2006/03/18 TDB

            const double t1 = 53808.0; // MJD at start of interval
            const double t2 = 53812.0; // MJD at end   of interval

            double[] Cx_moon =         // Coefficients for x-coordinate N_coeff
            {
                -0.383089044877016277e+06,  0.218158411754834669e+05,  0.179067292901463843e+05,
                -0.836928063411765777e+02, -0.628266733052023696e+02, -0.459274434235101225e+00,
                0.491167202819885532e-01,   0.770804039287614762e-03, -0.125935992206166816e-03,
                0.500271026610991370e-05,   0.107044869185752331e-05,  0.172472464343636242e-08,
                -0.269667589576924680e-08
            };
            double[] Cy_moon =     // Coefficients for y-coordinate N_coeff
            {
                -0.379891721705081436e+05, -0.143611643157166138e+06,  0.187126702787245881e+04,
                0.112734362473135207e+04,   0.932891213817359177e+00, -0.191932684130578513e+01,
                -0.266517663331897990e-01,  0.104558913448630337e-02, -0.359077689123857890e-04,
                -0.123405162037249834e-04,  0.180479239596339495e-06,  0.525522632333670539e-07,
                0.543313967008773005e-09
            };
            double[] Cz_moon =     // Coefficients for z-coordinate
            {
                -0.178496690739133737e+05, -0.788257550331743259e+05,  0.880684692614081882e+03,
                0.618395886330471512e+03,   0.103331218594995988e+01, -0.104949867328178592e+01,
                -0.150337371962561087e-01,  0.569056416308259317e-03, -0.186297523286550968e-04,
                -0.680012420653791955e-05,  0.902057208454410917e-07,  0.287891446432139173e-07,
                0.319822827699973363e-09
            };

            // Chebyshev approximation of lunar coordinates

            MultiDimChebshevFitter MoonCheb = new MultiDimChebshevFitter(t1, t2,              // Order and interval
                                                                         new Vector(Cx_moon), // Coefficients
                                                                         new Vector(Cy_moon),
                                                                         new Vector(Cz_moon));

            // Variables
            int    i;
            double Mjd0, Mjd_TT;
            Vector r = new Vector(3);

            // Epoch
            Mjd0 = DateUtil.DateToMjd(2006, 03, 14, 00, 00, 0.0);

            // Output
            var info = "Exercise 3-2: Lunar Ephemerides " + "\r\n";

            info += " Moon position from low precision analytical theory" + "\r\n";
            info += " Date [TT]                 " + " Position [km] " + "\r\n";
            Console.WriteLine(info);
            for (i = 0; i <= N_Step; i++)
            {
                Mjd_TT = Mjd0 + i * Step;
                r      = CelestialUtil.Moon(Mjd_TT) / 1000.0;

                info = " " + DateUtil.MjdToDateTimeString(Mjd_TT) + " " + r;

                Console.WriteLine(info);
            }

            info  = " Moon position from DE405" + "\r\n";
            info += " Date [TT]                 " + " Position [km] " + "\r\n";
            Console.WriteLine(info);
            for (i = 0; i <= N_Step; i++)
            {
                Mjd_TT = Mjd0 + i * Step;
                r      = MoonCheb.Fit(Mjd_TT);
                info   = " " + DateUtil.MjdToDateTimeString(Mjd_TT) + " " + r;
                Console.WriteLine(info);
            }

            Console.ReadKey();
        }